Combustor with baffle

ABSTRACT

A turbine engine and associated methods for a combustor as shown and described. The turbine engine including a combustor with a combustor liner having dilution openings and a geometry that changes along an axial direction. The combustor further having a baffle surrounding a combustor liner defining a combustion chamber of the combustor. A method for controlling nitrogen oxides within the combustor, including injecting compressed air into the annular combustion chamber through any of the dilution openings described herein.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims priority to U.S. Provisional Application SerialNo. 63/298,749, filed Jan. 12, 2022, the contents of which areincorporated herein by reference.

TECHNICAL FIELD

The present subject matter relates generally to a combustor for a gasturbine engine, and, more specifically to a combustor for burninglighter than air fuel, alone, or in a mixture.

BACKGROUND

Turbine engines are driven by a flow of combustion gases passing throughthe engine to rotate a multitude of turbine blades, which, in turn,rotate a compressor to provide compress air to the combustor forcombustion. A combustor can be provided within the turbine engine and isfluidly coupled with a turbine into which the combusted gases flow.

The use of hydrocarbon fuels in the combustor of a turbine engine isknown. Generally, air and fuel are fed to a combustion chamber, the airand fuel are mixed, and then the fuel is burned in the presence of theair to produce hot gas. The hot gas is then fed to a turbine where itcools and expands to produce power. By-products of the fuel combustiontypically include environmentally unwanted byproducts, such as nitrogenoxide and nitrogen dioxide (collectively called NO_(x)), carbon monoxide(CO), unburned hydrocarbons (UHC) (e.g., methane and volatile organiccompounds that contribute to the formation of atmospheric ozone), andother oxides, including oxides of sulfur (e.g., SO₂ and SO₃).

Varieties of fuel for use in combustion turbine engines are beingexplored. Hydrogen or hydrogen mixed with another element or compoundcan be used for combustion, however hydrogen or a hydrogen mixed fuelcan result in a higher flame temperature than traditional fuels. Thatis, hydrogen or a hydrogen mixed fuel typically has a wider flammablerange and a faster burning velocity than traditional fuels suchpetroleum-based fuels, or petroleum and synthetic fuel blends.

Standards stemming from air pollution concerns worldwide regulate theemission of NO_(x), UHC, and CO generated as a result of the turbineengine operation. In particular, NO_(x) is formed within the combustoras a result of high combustor flame temperatures during operation. It isdesirable to decrease NO_(x) emissions while still maintaining desirableefficiencies by regulating the profile and or pattern within thecombustor.

BRIEF DESCRIPTION OF THE DRAWINGS

In the drawings:

FIG. 1 is a schematic of a turbine engine.

FIG. 2 depicts a cross-section view along line II-II of FIG. 1 of acombustion section of the turbine engine.

FIG. 3 is a cross-sectional view along line III-III of FIG. 2 of acombustor in the combustion section with a baffle defining a shelteredzone.

FIG. 4 is a schematic of a portion of the combustor taken along line A-Aof FIG. 3 .

FIG. 5 is a schematic of a portion of the combustor taken along line B-Bof FIG. 3 .

FIG. 6 is a schematic of a portion of the combustor taken along line E-Eof FIG. 3 .

FIG. 7 is a variation of the portion of the combustor taken along lineA-A of FIG. 3 .

FIG. 8 is the portion of the combustor taken along line B-B of FIG. 3 .

FIG. 9 is the portion of the combustor taken along line E-E of FIG. 3 .

FIG. 10 is an enlarged view of a portion of the combustor of FIG. 3illustrating a set of acoustic dampers located within the shelteredzone.

FIG. 11 is schematic of a portion of a combustor in a cross-sectionalview, similar to the combustor of FIG. 3 , with a baffle defining asheltered zone according to another aspect of the disclosure herein.

FIG. 12 is schematic of a portion of a combustor in a cross-sectionalview, similar to the combustor of FIG. 3 , with a baffle defining asheltered zone according to yet another aspect of the disclosure herein.

FIG. 13 is schematic of a portion of a combustor in a cross-sectionalview, similar to the combustor of FIG. 3 , with a baffle defining asheltered zone according to yet another aspect of the disclosure herein.

DETAILED DESCRIPTION

Aspects of the disclosure described herein are directed to a combustor,and in particular a combustor with a baffle defining a sheltered zoneproximate a fuel injector of the combustor. Further the combustor asdescribed herein can include a combustion chamber having a primary zonewith a transitional cross-section from a can to an annular profile. Forpurposes of illustration, the present disclosure will be described withrespect to a turbine engine. It will be understood, however, thataspects of the disclosure described herein are not so limited and that acombustor as described herein can be implemented in engines, includingbut not limited to turbojet, turboprop, turboshaft, and turbofanengines. Aspects of the disclosure discussed herein may have generalapplicability within non-aircraft engines having a combustor, such asother mobile applications and non-mobile industrial, commercial, andresidential applications.

The word “exemplary” is used herein to mean “serving as an example,instance, or illustration.” Any implementation described herein as“exemplary” is not necessarily to be construed as preferred oradvantageous over other implementations. Additionally, unlessspecifically identified otherwise, all embodiments described hereinshould be considered exemplary.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine or vehicle, and refer to the normal operational attitudeof the gas turbine engine or vehicle. For example, with regard to a gasturbine engine, forward refers to a position closer to an engine inletand aft refers to a position closer to an engine nozzle or exhaust.

As used herein, the term “upstream” refers to a direction that isopposite the fluid flow direction, and the term “downstream” refers to adirection that is in the same direction as the fluid flow. The term“fore” or “forward” means in front of something and “aft” or “rearward”means behind something. For example, when used in terms of fluid flow,fore/forward can mean upstream and aft/rearward can mean downstream.

The term “fluid” may be a gas or a liquid. The term “fluidcommunication” means that a fluid is capable of making the connectionbetween the areas specified.

Additionally, as used herein, the terms “radial” or “radially” refer toa direction away from a common center. For example, in the overallcontext of a turbine engine, radial refers to a direction along a rayextending between a center longitudinal axis of the engine and an outerengine circumference.

All directional references (e.g., radial, axial, proximal, distal,upper, lower, upward, downward, left, right, lateral, front, back, top,bottom, above, below, vertical, horizontal, clockwise, counterclockwise,upstream, downstream, forward, aft, etc.) are only used foridentification purposes to aid the reader’s understanding of the presentdisclosure, and do not create limitations, particularly as to theposition, orientation, or use of aspects of the disclosure describedherein. Connection references (e.g., attached, coupled, connected, andjoined) are to be construed broadly and can include intermediatestructural elements between a collection of elements and relativemovement between elements unless otherwise indicated. As such,connection references do not necessarily infer that two elements aredirectly connected and in fixed relation to one another. The exemplarydrawings are for purposes of illustration only the dimensions,positions, order and relative sizes reflected in the drawings attachedhereto can vary.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise. Furthermore, as used herein, theterm “set” or a “set” of elements can be any number of elements,including only one.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about”, “approximately”, “generally”, and “substantially”, arenot to be limited to the precise value specified. In at least someinstances, the approximating language may correspond to the precision ofan instrument for measuring the value, or the precision of the methodsor machines for constructing or manufacturing the components and/orsystems. In at least some instances, the approximating language maycorrespond to the precision of an instrument for measuring the value, orthe precision of the methods or machines for constructing ormanufacturing the components and/or systems. For example, theapproximating language may refer to being within a 1, 2, 4, 5, 10, 15,or 20 percent margin in either individual values, range(s) of valuesand/or endpoints defining range(s) of values. Here and throughout thespecification and claims, range limitations are combined andinterchanged, such ranges are identified and include all the sub-rangescontained therein unless context or language indicates otherwise. Forexample, all ranges disclosed herein are inclusive of the endpoints, andthe endpoints are independently combinable with each other.

“Proximate” as used herein is not limiting, rather a descriptor forlocating parts described herein. Further, the term “proximate” meansnearer or closer to the part recited than the following part. Forexample, a first hole proximate a wall, the first hole located upstreamfrom a second hole means that the first hole is closer to the wall thanthe first hole is to the second hole.

FIG. 1 is a schematic view of a turbine engine 10. As a non-limitingexample, the turbine engine 10 can be used within an aircraft. Theturbine engine 10 can include, at least, a compressor section 12, acombustion section 14, and a turbine section 16. A drive shaft 18rotationally couples the compressor and turbine sections 12, 16, suchthat rotation of one affects the rotation of the other, and defines arotational axis or centerline 20 for the turbine engine 10.

The compressor section 12 can include a low-pressure (LP) compressor 22,and a high-pressure (HP) compressor 24 serially fluidly coupled to oneanother. The turbine section 16 can include an LP turbine 26, and an HPturbine 28 serially fluidly coupled to one another. The drive shaft 18can operatively couple the LP compressor 22, the HP compressor 24, theLP turbine 26 and the HP turbine 28 together. Alternatively, the driveshaft 18 can include an LP drive shaft (not illustrated) and an HP driveshaft (not illustrated). The LP drive shaft can couple the LP compressor22 to the LP turbine 26, and the HP drive shaft can couple the HPcompressor 24 to the HP turbine 28. An LP spool can be defined as thecombination of the LP compressor 22, the LP turbine 26, and the LP driveshaft such that the rotation of the LP turbine 26 can apply a drivingforce to the LP drive shaft, which in turn can rotate the LP compressor22. An HP spool can be defined as the combination of the HP compressor24, the HP turbine 28, and the HP drive shaft such that the rotation ofthe HP turbine 28 can apply a driving force to the HP drive shaft whichin turn can rotate the HP compressor 24.

The compressor section 12 can include a plurality of axially spacedstages. Each stage includes a set of circumferentially-spaced rotatingblades and a set of circumferentially-spaced stationary vanes. Thecompressor blades for a stage of the compressor section 12 can bemounted to a disk, which is mounted to the drive shaft 18. Each set ofblades for a given stage can have its own disk. The vanes of thecompressor section 12 can be mounted to a casing which can extendcircumferentially about the turbine engine 10. It will be appreciatedthat the representation of the compressor section 12 is merely schematicand that there can be any number of stages. Further, it is contemplated,that there can be any other number of components within the compressorsection 12.

Similar to the compressor section 12, the turbine section 16 can includea plurality of axially spaced stages, with each stage having a set ofcircumferentially-spaced, rotating blades and a set ofcircumferentially-spaced, stationary vanes. The turbine blades for astage of the turbine section 16 can be mounted to a disk which ismounted to the drive shaft 18. Each set of blades for a given stage canhave its own disk. The vanes of the turbine section can be mounted tothe casing in a circumferential manner. It is noted that there can beany number of blades, vanes and turbine stages as the illustratedturbine section is merely a schematic representation. Further, it iscontemplated, that there can be any other number of components withinthe turbine section 16.

The combustion section 14 can be provided serially between thecompressor section 12 and the turbine section 16. The combustion section14 can be fluidly coupled to at least a portion of the compressorsection 12 and the turbine section 16 such that the combustion section14 at least partially fluidly couples the compressor section 12 to theturbine section 16. As a non-limiting example, the combustion section 14can be fluidly coupled to the HP compressor 24 at an upstream end of thecombustion section 14 and to the HP turbine 28 at a downstream end ofthe combustion section 14.

During operation of the turbine engine 10, ambient or atmospheric air isdrawn into the compressor section 12 via a fan (not illustrated)upstream of the compressor section 12, where the air is compresseddefining a pressurized air. The pressurized air can then flow into thecombustion section 14 where the pressurized air is mixed with fuel andignited, thereby generating combustion gases. Some work is extractedfrom these combustion gases by the HP turbine 28, which drives the HPcompressor 24. The combustion gases are discharged into the LP turbine26, which extracts additional work to drive the LP compressor 22, andthe exhaust gas is ultimately discharged from the turbine engine 10 viaan exhaust section (not illustrated) downstream of the turbine section16. The driving of the LP turbine 26 drives the LP spool to rotate thefan (not illustrated) and the LP compressor 22. The pressurized airflowand the combustion gases can together define a working airflow thatflows through the fan, compressor section 12, combustion section 14, andturbine section 16 of the turbine engine 10.

FIG. 2 depicts a cross-sectional view of the combustion section 14 alongline II-II of FIG. 1 . The combustion section 14 can include a set offuel injectors 30 annularly arranged about the centerline 20 of theturbine engine 10. A combustor 34 is fluidly connected to the set offuel injectors 30 to define at least a portion of a set of fuel cups 32.The set of fuel cups 32 can include rich cups, lean cups, or acombination of both rich and lean cups annularly provided about theengine centerline 20. It should be appreciated that the annulararrangement of fuel injectors 30 can be one or multiple fuel injectors30 and one or more of the fuel injectors 30 can have differentcharacteristics.

A combustor liner 38 including an outer combustor liner 40 and an innercombustor liner 42 concentric with respect to each other and annularabout the engine centerline 20 defines the combustor 34. The combustorliner 38 also further defines the set of fuel cups 32. A dome wall 44together with the combustor liner 38 can define a combustion chamber 46of the combustor 34 annular about the engine centerline 20. The set offuel cups 32 can be fluidly coupled to the combustion chamber 46.

The combustor 34 can have a can, can-annular, or annular arrangementdepending on the type of engine in which the combustor 34 is located.The combustor liner 38 can have a varying geometry as further describedherein. The combustor 34 can be fully encased by a casing 36. Acompressed air passageway 48 can be defined at least in part by both thecombustor liner 38 and the casing 36.

FIG. 3 depicts a cross-sectional view taken along line III-III of FIG. 2illustrating the combustion section 14. A dome assembly 50 can house thefuel injector 30. The fuel injector 30 can be fluidly coupled to a fuelinlet 52 via a fuel passageway 54 that can be adapted to receive a flowof fuel (F). Compressed air (C) can be provided to the combustionsection 14 from the compressor section 12 via the compressed airpassageway 48. The fuel injector 30 can terminate in a dome inlet 56 todefine the fuel cup 32. The dome inlet 56 can further define along withthe combustor liner 38, at least a portion of a primary zone 58 in thecombustion chamber 46. A swirler 60 can be fluidly coupled to the fuelinjector 30. A first set of dilution openings 62 can be provided in thecombustor liner 38 for connecting the compressed air passageway 48 andthe combustion chamber 46. At least one igniter 64 can be coupled to thecombustor liner 38.

A baffle 70 can be mounted to an outer surface 66 of the combustor liner38 surrounding the primary zone 58. The baffle 70 can be spaced from thecombustor liner 38 to define a sheltered zone 72 with a shelter zoneinlet 74. The baffle 70 can direct a specific amount of compressed air(C) for usage along the combustor liner 38. A second set of dilutionopenings 76 can be provided in the combustor liner 38 for fluidlyconnecting the sheltered zone 72 to the primary zone 58 of thecombustion chamber 46. A third set of dilution openings 78 can beprovided in the combustor liner 38 downstream of the second set ofdilution openings 76 for connecting a different portion of the shelteredzone 72 to the primary zone 58 of the combustion chamber 46. The first,second, and third sets of dilution openings can be round or shapedholes, slots, or annular gaps and designed to form radial or angled airjets.

During operation, compressed air (C) can be fed into the fuel injector30 and mixed with fuel (F) to define a fuel/air mixture. The mixture canbe ignited within the combustion chamber 46 by the at least one igniter64 to generate combustion gas (G). The swirler 60 can swirl incomingcompressed air (C) with fuel (F) entering the fuel cup 32 to provide ahomogeneous mixture of air and fuel entering the combustion chamber 46via the dome inlet 56.

Further, compressed air (C) can be fed into the sheltered zone 72 viathe shelter zone inlet 74. Compressed air (C) can be utilized asdilution jets provided through both the second and third set of dilutionopenings 76, 78. The primary zone 58 between the second and third set ofdilution openings 76, 78 can define an extension of the at least onefuel cup 32. The baffle 70 provides the sheltered zone 72 for the fuelcup 32 such that the primary zone 58 operates similarly to an individualcan combustor arrangement. The baffle 70 allows for stability of asingle one of the fuel cups 32 by minimizing cup to cup interaction.

Compressed air (C) can additionally enter the combustion chamber 46 viathe first set of dilution openings 62 to provide a dilution flow (D)within the combustion chamber 46. The combustion gas (G) can be mixedusing the dilution flow (D) or simply controlled by the dilution flow(D) to move through a combustor outlet 80 and exit into the turbinesection 16.

FIGS. 4-6 illustrate varying cross-sections of the combustor 34 atvarying locations moving axially from a first end 43 at the dome wall 44toward a second end 45 at the combustor outlet 80. As illustrated inFIGS. 4-6 , the combustor liner 38 has a transitional geometry. Thecombustor liner 38 transitions from a defined cup geometry 100 (FIG. 4 )proximate the set of fuel injectors 30 to an annular geometry 102 (FIG.6 ) axially downstream the set of fuel injectors 30. The defined cupgeometry meaning the combustor liner 38 is rounded in a cylindricalshape to define distinguishable cups. More specifically the geometry ofthe combustor liner 38 transitions from a modified can shape 104 (FIG. 4) to a can-annular shape 106 (FIG. 5 ) to an annular shape 108 (FIG. 6).

FIG. 4 is a cross-sectional view facing multiple fuel injectors 30 alongline A-A of FIG. 3 looking toward the first end 43 and illustrating themodified can shape 104. The second set of dilution openings 76 caninclude multiple dilution openings annular about the fuel cup 32. Thecombustor liner 38 can be shaped to define distinct fuel cups, a firstfuel cup 32 a, a second fuel cup 32 b, and a third fuel cup 32 c fluidlyconnected to each other by a spacing of the inner and outer combustorliners 40, 42 defining circumferentially spaced openings 82. Themodified can shape 104 is defined as having the openings 82 that aresmall enough to define the plurality of discrete,circumferentially-spaced, fuel cups 32 a, 32 b, 32 c while stillenabling a fluid connection between the distinct fuel cups 32 a, 32 b,32 c. The baffle 70 (FIG. 3 ) described herein enables proper feeding ofcompressed air (C) to the second and third sets of dilution openings 76,78 in the sheltered zone 72 (FIG. 3 ) to produce operation similar tothat of an individual can combustor while geometrically shaped with themodified can shape 104.

FIG. 5 is a cross-sectional view facing multiple fuel injectors 30 alongline B-B of FIG. 3 just downstream from the third set of dilutionopenings 78. The circumferentially spaced openings 82 can graduallywiden as the combustor liner 38 extends downstream toward the second end45 between line A-A and line B-B of FIG. 3 . It can be appreciated thatthe combustor liner 38 at line B-B is shaped such that the distinct richfuel cups 32 a, 32 b, 32 c are still distinguishable but more fluidlydynamic with respect to each other. In this manner, the geometry of thecombustor liner 38 transitions to define a combustor 34 with a modifiedcan shape 104 at line A-A to a can-annular shape 106 at line B-B.

FIG. 6 is a cross-sectional view facing multiple fuel injectors 30 alongline E-E of FIG. 3 just downstream from line B-B. The circumferentiallyspaced openings 82 (FIGS. 4 and 5 ) are no longer distinguishable. Itcan be appreciated that the combustor liner 38 at line E-E is shapedsuch that the distinct fuel cups 32 a, 32 b, 32 c (FIGS. 4 and 5 ) areno longer distinguishable and the combustion chamber 46 is annular aboutthe engine centerline 20 (FIG. 2 ). The outer and inner combustor liners40, 42 at line E-E are circular in shape and annular about the enginecenterline 20 (FIG. 2 ). In this manner, the geometry of the combustorliner 38 transitions to define a combustor 34 with a can-annular shape106 at line B-B to an annular shape 108 at E-E.

FIGS. 7-9 illustrates varying cross-sectional schematics of thecombustor 34 according to another aspect of the disclosure herein.Moving axially from the first end 43 at the dome wall 44 toward thesecond end 45 at the combustor outlet 80, a combustor liner 138 has atransitional geometry. The combustor liner 138 is substantially similarto the combustor liner 38, therefore, like parts will be identified withlike numerals increased by 100. It should be understood that thedescription of the like parts of the combustor liner 38 applies to thecombustor liner 138 unless otherwise noted.

The geometry of the combustor liner 138 transitions from a can shape 110(FIG. 7 ) to a can-annular shape 106 (FIG. 8 ) to an annular shape 108(FIG. 9 ). The can shape 110 (FIG. 7 ) is defined as having no openingssuch that the combustor liner 138 defines a plurality of discrete,circumferentially-spaced, fuel cups 132 a, 132 b, 132 c without a fluidconnection between the distinct fuel cups 132 a, 132 b, 132 c at thefirst end 43. More specifically inner and outer combustor liners 140,142 of the combustor liner 138 meet to physically separate rich fuelcups 132 a, 132 b, 132 c at line A-A (FIG. 3 ). While at line B-B (FIG.3 ), the inner and outer combustor liners 140, 142 are separated todefine a set of spaced openings 182, in turn defining the can-annularshape 106 (FIG. 8 ). It can be appreciated that the combustor liner 138at line E-E (FIG. 3 ) is shaped such that the distinct fuel cups 132 a,132 b, 132 c (FIGS. 7 and 8 ) are no longer distinguishable, in turndefining the combustion chamber 146 in the annular shape 108 (FIG. 9 ).

FIG. 10 schematic of exemplary combustor 34 illustrating a variation ofthe baffle 70 according to an aspect of the disclosure herein. A baffle170 is substantially similar to the baffle 70, therefore, like partswill be identified with like numerals increased by 100. It should beunderstood that the description of the like parts of the baffle 70applies to the baffle 170 unless otherwise noted.

The baffle 170 can be spaced from the outer surface 66 of the combustorliner 38 to define a sheltered zone 172. The baffle 170 can include abody 184 extending axially from a mounting leg 186. The baffle 170 canbe mounted to the outer surface 66 of the combustor liner 38 at themounting leg 186. The mounting leg 186 can be mounted to the outersurface 66 between the second and third set of dilution openings 76, 78.The body 184 can be spaced from the outer surface 66 proximate the domewall 44 to define a shelter zone inlet 174 fluidly coupled to thesheltered zone 172.

The baffle 170 can divide the area around the combustor 34 into a totalpressure zone 188 and a static pressure zone 190. During operation thetotal pressure zone 188 feeds the second set of dilution openings 76,while static pressure zone 190 feeds the third set of dilution openings78. It is contemplated that any number of dilution openings are part ofthe first set of dilution openings 62, which can also be fed by thestatic pressure zone 190.

FIG. 11 is a schematic of exemplary combustor 34 illustrating avariation of the baffle 70 according to another aspect of the disclosureherein. A baffle 270 is substantially similar to the baffle 70,therefore, like parts will be identified with like numerals increased by200. It should be understood that the description of the like parts ofthe baffle 70 applies to the baffle 170 unless otherwise noted.

The baffle 270 can be spaced from the outer surface 66 of the combustorliner 38 to define a sheltered zone 272. The baffle 270 can include abody 284 extending axially from a mounting leg 286. The baffle 270 canbe mounted to the outer surface 66 of the combustor liner 38 at themounting leg 286. The body 284 can be spaced from the outer surface 66proximate the dome wall 44 to define a shelter zone inlet 274 fluidlycoupled to the sheltered zone 272. The mounting leg 286 can be mountedto the outer surface 66 downstream of the first set of dilution openings62.

The baffle 270 can divide the area around the combustor 34 into a totalpressure zone 288 and a static pressure zone 290. During operation thetotal pressure zone 288 feeds first, second, and third sets of dilutionopenings 62, 76, 78 described herein. Feeding the dilution openingsdescribed herein with total or static pressure enables tuning of anamount of pressure drop across the dilution openings. This amount ofpressure drop is directly related to an amount of penetration of thecompressed air (C) as a dilution jet in the combustion chamber 46. Thistuning has been demonstrated to impact emissions.

FIG. 12 is a schematic of exemplary combustor 34 illustrating avariation of the baffle 70 according to another aspect of the disclosureherein. A baffle 370 is substantially similar to the baffle 70,therefore, like parts will be identified with like numerals increased by300. It should be understood that the description of the like parts ofthe baffle 70 applies to the baffle 370 unless otherwise noted.

The baffle 370 can be spaced from the outer surface 66 of the combustorliner 38 to define a sheltered zone 372. The baffle 370 can include abody 384 extending axially between a mounting leg 386 and an inlet leg392. The baffle 370 can be mounted to the outer surface 66 of thecombustor liner 38 at the mounting leg 386. The inlet leg 392 can bespaced from the dome wall 44 to define a shelter zone inlet 374 fluidlycoupled to the sheltered zone 372. The baffle 370 can have asubstantially trapezoidal shape in cross-section where the shelter zoneinlet 374 is located along a base of the trapezoidal shape.

The baffle 370 can divide the area around the combustor 34 into a totalpressure zone 388 and a static pressure zone 390. During operation thetotal pressure zone 388 feeds second and third sets of dilution openings76, 78, while static pressure zone 390 feeds the first set of dilutionopenings 62. As previously described, feeding the dilution openings withtotal or static pressure enables tuning of an amount of pressure dropacross the dilution openings. While illustrated with respect to thebaffle 370, it should be understood that the inlet leg 392 describedherein can be associated with any of the baffles 70, 170, 270 describedherein.

Turning to FIG. 13 , an enlarged view of the baffle 370 according toanother aspect of the disclosure herein is illustrated. At least oneacoustic damper 394 can be provided in the sheltered zone 372 and extendradially from the outer surface 66 to the body 384. The at least oneacoustic damper 394 can be multiple dampers, a first acoustic damper 394a placed between the second and third set of dilution openings 76, 78and a second acoustic damper 394 b placed downstream from the third setof dilution openings 78 and upstream from the mounting leg 386.

The baffle 370 provides a coupling between the acoustic damper 394 and asource of acoustic pressure caused by acoustic dynamics associated withchamber volume, heat release variation during combustion, and swirlerflow vorticity. The acoustic damper 394 can be provided to reducepressure fluctuations at specific frequencies. The acoustic damper 394can dampen an associated amplitude of the specific frequencies. Theacoustic damper 394 can also be called a Helmholtz resonator. The sizeof the damper is directly related to the acoustic frequency. Theacoustic damper 394 is especially beneficial in combustors withrelatively high amplitudes associated with the specific frequencies.While illustrated with respect to the baffle 370, it should beunderstood that the acoustic dampers described herein can be located inany of the baffles 70, 170, 270 described herein

Any combination of the arrangements with regard to the baffles describedherein are contemplated. The rich cup and lean cup arrangements can bein any form described herein.

A method for controlling nitrogen oxides, or NO_(x) present incombustion gases (G) within the combustor, includes injecting compressedair (C) into the combustion chamber through any of the dilution openingsdescribed herein. The method can further includes swirling thecompressed air (C) with fuel (F) before injection into the combustionchamber. Mixing the compressed air (C) with the fuel (F) can occur inthe primary zone. The method can further include transitioning the flowof combustion gasses (G) through the primary zone 58 to the combustoroutlet 80 from distinguishable cups 32 to a single annular geometry 102.The transition in combustor liner geometry can also be utilized forcontrolling a flame of the fuel/air mixture in the primary zone 58.

Benefits associated with the baffle and transitional geometry of thecombustor liner and methods described herein are a reduction and/orelimination of CO emissions. Further, the transitional geometry of thecombustor liner contributes to flame control. The baffle and positioningof said baffle described herein also contributes to controlling theflame produced by H₂ fuel to achieve lower NO_(x), lower dynamics andbetter component life.

While described with respect to a turbine engine, it should beappreciated that the combustor as described herein can be for any enginewith a having a combustor that emits NO_(x). It should be appreciatedthat application of aspects of the disclosure discussed herein areapplicable to engines with propeller sections or fan and boostersections along with turbojets and turbo engines as well.

To the extent not already described, the different features andstructures of the various embodiments can be used in combination, or insubstitution with each other as desired. That one feature is notillustrated in all of the embodiments is not meant to be construed thatit cannot be so illustrated, but is done for brevity of description.Thus, the various features of the different embodiments can be mixed andmatched as desired to form new embodiments, whether or not the newembodiments are expressly described. All combinations or permutations offeatures described herein are covered by this disclosure.

This written description uses examples to describe aspects of thedisclosure described herein, including the best mode, and also to enableany person skilled in the art to practice aspects of the disclosure,including making and using any devices or systems and performing anyincorporated methods. The patentable scope of aspects of the disclosureis defined by the claims, and may include other examples that occur tothose skilled in the art. Such other examples are intended to be withinthe scope of the claims if they have structural elements that do notdiffer from the literal language of the claims, or if they includeequivalent structural elements with insubstantial differences from theliteral languages of the claims.

Further aspects are provided by the subject matter of the followingclauses:

A gas turbine engine comprising a compressor section, combustionsection, and turbine section in serial flow arrangement, the combustionsection comprising an annular combustion chamber having a first end anda second end, wherein the second end is spaced from the first end; aplurality of fuel cups located at the first end and annularly arrangedwithin the annular combustion chamber, wherein at least one of theplurality of fuel cups has a defined cup geometry; and a combustor linerat least partially defining the annular combustion chamber and having avarying cross-section extending in an axial direction from the first endtoward the second end, the varying cross-section defining at the firstend at least a portion of the defined cup geometry and at the second enda single annular geometry.

The gas turbine engine of any preceding clause wherein the defined cupgeometry is a modified can shape defining a plurality of discrete,circumferentially-spaced, fuel cups fluidly connected to each otherproximate the first end

The gas turbine engine of any preceding clause wherein the defined cupgeometry is a can shape defining a plurality of discrete,circumferentially-spaced, fuel cups fluidly separate from each otherproximate the first end.

The gas turbine engine of any preceding clause wherein the combustorliner is shaped to define a can-annular shape downstream from the firstend and upstream from the second end.

The gas turbine engine of any preceding clause, further comprising atleast one set of dilution openings in the combustor liner.

The gas turbine engine of any preceding clause, further comprising abaffle mounted to an outer surface of the combustor liner to define asheltered zone.

The gas turbine engine of any preceding clause wherein the at least oneset of dilution openings fluidly connects the sheltered zone to theannular combustion chamber.

The gas turbine engine of any preceding clause wherein the at least oneset of dilution openings is multiple sets of dilution openings axiallyspaced from each other along the combustor liner.

The gas turbine engine of any preceding clause wherein the at least oneset of dilution openings is at least two sets of dilution openingsaxially spaced from each other along the combustor liner.

The gas turbine engine of any preceding clause wherein the baffle ismounted to the outer surface between the at least two sets of dilutionopenings.

The gas turbine engine of any preceding clause wherein at least one setof dilution openings fluidly connect the sheltered zone to the annularcombustion chamber.

The gas turbine engine of any preceding clause wherein the bafflecomprises a body extending axially between a mounting leg and an inletleg and the baffle is mounted to the outer surface of the combustorliner at the mounting leg and the inlet leg is spaced from the first endto define a shelter zone inlet fluidly coupled to the sheltered zone.

The gas turbine engine of any preceding clause wherein the bafflecomprises an axially extending body.

The gas turbine engine of any preceding clause wherein a mounting legextends from the body.

The gas turbine engine of any preceding clause wherein the mounting legis mounted to the combustor liner and the body is spaced from thecombustor liner to define a sheltered zone.

The gas turbine engine of any preceding clause wherein the baffledivides an area around the combustor into a total pressure zone and astatic pressure zone wherein the total pressure zone and the shelteredzone are on in the same.

The gas turbine engine of any preceding clause wherein at least one setof dilution openings is fed by the total pressure zone and another setof dilution openings is fed by the static pressure zone.

The gas turbine engine of any preceding clause wherein all of thedilution openings are fed by the total pressure zone.

The gas turbine engine of any preceding clause wherein an inlet legextends from the body to further define the sheltered zone and a shelterzone inlet.

A combustor comprising an annular combustion chamber having a first endand a second end, wherein the second end is spaced from the first end; aplurality of fuel cups located at the first end and annularly arrangedwithin the annular combustion chamber, wherein at least some of theplurality of fuel cups have a defined cup geometry; and a combustorliner at least partially defining the annular combustion chamber andhaving a varying cross-section extending in a direction downstream fromthe first end toward the second end, the varying cross-section definingat the first end at least a portion of the defined cup geometry for thefuel cups and at the second end a single annular geometry.

The combustor of any preceding clause wherein the defined cup geometryis a modified can shape defining a plurality of discrete,circumferentially-spaced, fuel cups fluidly connected to each otherproximate the first end.

The combustor of any preceding clause wherein the defined cup geometryis a can shape defining a plurality of discrete,circumferentially-spaced, fuel cups fluidly separate from each otherproximate the first end.

The combustor of any preceding clause wherein the combustor liner isshaped to define a can-annular shape downstream from the defined cupgeometry and upstream from the single annular geometry.

The combustor of any preceding clause, further comprising at least oneset of dilution openings in the combustor liner and a baffle mounted toan outer surface of the combustor liner to define a sheltered zone, atleast one set of the at least one set of dilution openings fluidlyconnecting the sheltered zone to the annular combustion chamber.

A method for controlling nitrogen oxides present within a combustor of aturbine engine, the method comprising injecting a fuel and a compressedair mixture into a combustion chamber of the combustor through a fuelinjector; mixing the compressed air and the fuel to define a fuel/airmixture; igniting the fuel/air mixture in a primary zone of thecombustion chamber to define a flame and to generate combustion gasses;flowing the combustion gasses through the combustion chamber from aplurality of circumferentially spaced, discrete cups to a single annulargeometry; and exhausting the combustion gasses at a combustor outlet.

The method of any preceding clause, further comprising controlling aflame of the fuel/air mixture in the primary zone by feeding compressedair into a sheltered zone defined by a baffle mounted to an outersurface of the combustor.

The method of any preceding clause, further comprising reducing pressurefluctuations at specific frequencies with an acoustic damper locatedwithin the sheltered zone.

The method of any preceding clause, further comprising feeding a set ofdilution openings from the sheltered zone.

The method of any preceding clause, further comprising feeding a set ofdilution openings from a static pressure zone.

The method of any preceding clause, further comprising feeding a set ofdilution openings from a total pressure zone.

What is claimed is:
 1. A gas turbine engine comprising: a compressorsection, combustion section, and turbine section in serial flowarrangement, the combustion section comprising: an annular combustionchamber having a first end and a second end, wherein the second end isspaced from the first end; a plurality of fuel cups located at the firstend and annularly arranged within the annular combustion chamber,wherein at least one of the plurality of fuel cups has a defined cupgeometry; and a combustor liner at least partially defining the annularcombustion chamber and having a varying cross-section extending in anaxial direction from the first end toward the second end, the varyingcross-section defining at the first end at least a portion of thedefined cup geometry and at the second end a single annular geometry. 2.The gas turbine engine of claim 1 wherein the defined cup geometry is amodified can shape defining a plurality of discrete,circumferentially-spaced, fuel cups fluidly connected to each otherproximate the first end.
 3. The gas turbine engine of claim 1 whereinthe defined cup geometry is a can shape defining a plurality ofdiscrete, circumferentially-spaced, fuel cups fluidly separate from eachother proximate the first end.
 4. The gas turbine engine of claim 1wherein the combustor liner is shaped to define a can-annular shapedownstream from the first end and upstream from the second end.
 5. Thegas turbine engine of claim 1, further comprising at least one set ofdilution openings in the combustor liner.
 6. The gas turbine engine ofclaim 5, further comprising a baffle mounted to an outer surface of thecombustor liner to define a sheltered zone.
 7. The gas turbine engine ofclaim 6 wherein the at least one set of dilution openings fluidlyconnects the sheltered zone to the annular combustion chamber.
 8. Thegas turbine engine of claim 7 wherein the at least one set of dilutionopenings is multiple sets of dilution openings axially spaced from eachother along the combustor liner.
 9. The gas turbine engine of claim 6wherein the at least one set of dilution openings is at least two setsof dilution openings axially spaced from each other along the combustorliner.
 10. The gas turbine engine of claim 9 wherein the baffle ismounted to the outer surface between the at least two sets of dilutionopenings.
 11. The gas turbine engine of claim 10 wherein at least oneset of dilution openings fluidly connect the sheltered zone to theannular combustion chamber.
 12. The gas turbine engine of claim 6wherein the baffle comprises a body extending axially between a mountingleg and an inlet leg and the baffle is mounted to the outer surface ofthe combustor liner at the mounting leg and the inlet leg is spaced fromthe first end to define a shelter zone inlet fluidly coupled to thesheltered zone.
 13. A combustor comprising: an annular combustionchamber having a first end and a second end, wherein the second end isspaced from the first end; a plurality of fuel cups located at the firstend and annularly arranged within the annular combustion chamber,wherein at least some of the plurality of fuel cups have a defined cupgeometry; and a combustor liner at least partially defining the annularcombustion chamber and having a varying cross-section extending in adirection downstream from the first end toward the second end, thevarying cross-section defining at the first end at least a portion ofthe defined cup geometry for the fuel cups and at the second end asingle annular geometry.
 14. The combustor of claim 13 wherein thedefined cup geometry is a modified can shape defining a plurality ofdiscrete, circumferentially-spaced, fuel cups fluidly connected to eachother proximate the first end.
 15. The combustor of claim 13 wherein thedefined cup geometry is a can shape defining a plurality of discrete,circumferentially-spaced, fuel cups fluidly separate from each otherproximate the first end.
 16. The combustor of claim 13 wherein thecombustor liner is shaped to define a can-annular shape downstream fromthe defined cup geometry and upstream from the single annular geometry.17. The combustor of claim 13, further comprising at least one set ofdilution openings in the combustor liner and a baffle mounted to anouter surface of the combustor liner to define a sheltered zone, atleast one set of the at least one set of dilution openings fluidlyconnecting the sheltered zone to the annular combustion chamber.
 18. Amethod for controlling nitrogen oxides present within a combustor of aturbine engine, the method comprising: injecting a fuel and a compressedair mixture into a combustion chamber of the combustor through a fuelinjector; mixing the compressed air and the fuel to define a fuel/airmixture; igniting the fuel/air mixture in a primary zone of thecombustion chamber to define a flame and to generate combustion gasses;flowing the combustion gasses through the combustion chamber from aplurality of circumferentially spaced, discrete cups to a single annulargeometry; and exhausting the combustion gasses at a combustor outlet.19. The method of claim 18, further comprising controlling a flame ofthe fuel/air mixture in the primary zone by feeding compressed air intoa sheltered zone defined by a baffle mounted to an outer surface of thecombustor.
 20. The method of claim 19, further comprising reducingpressure fluctuations at specific frequencies with an acoustic damperlocated within the sheltered zone.